In general, a structure is said to be failure-mitigating or, more specifically, fail safe when there are several possible paths along which load can be reacted. In particular, a failure-mitigating structure may be made up of two longitudinal metal components assembled with one another to act as a reinforced frame, known as the “mainframe” of an aircraft fuselage. Because of the extreme magnitude of the loads applied, and because of the difficulties connected with manufacture, these frames are generally made of metal.
For such a mainframe to be certified both of its two components have to have mechanical integrity at 150% of the maximum possible loads encountered in the life of the aircraft (loads known as “ultimate loads”). When one of the two components is assumed to have broken, the mechanical integrity at 100% of the maximum load applied in the life of the aircraft (loads known as “limit loads”) has to be demonstrated.
Because fuselage mainframes are made of metal, the main criterion governing the sizing of these frames is generally their ability to tolerate damage. This is because certification demands that the largest crack that has remained undetected in a first inspection must not be able to grow to the critical size—defined as being able to destroy the structure—during the interval of time separating the first inspection from the next inspection.
In order to measure the damage tolerance of a fuselage mainframe, it is agreed practice to use a crack spread model that allows the size of the crack or cracks to be assessed as a function of the number of flights made. The initial conditions generally considered for a structure of the failsafe frame type generally consist in considering cracks of different sizes on each of the components of the failsafe frame. The cracks are formed on the two assembled components of the frame and then spread at rates that are proportional to the initial size of the cracks. When a crack has reached the critical rupture size, the corresponding component is broken and the other component finds itself overloaded because of the redistribution of load from the broken frame into the other frame, and into the skin of the fuselage. The order of magnitude of the overload experienced by the remaining frame is around 80%. This therefore corresponds to what is known as an overall redistribution of load. The spread of the crack through the unbroken frame is then very rapid. The interval between two inspections has therefore to be determined by the component sizing criterion.